This invention relates to spacecraft structures, and, in particular, relates to spacecraft structures used to hold multiple pieces of equipment in a rigid manner.
Primary considerations for such structural systems are (1) minimum weight, (2) highest possible structural rigidity, and (3) maximum thermal stability. A space-borne structural system must also withstand the rigors of launch and orbital environments without exhibiting distortions due to launch loads, gravity release, or exposure to a vacuum environment. Such a high precision structural system must also be capable of maintaining the alignment of its components to within its requirements while experiencing the induced environment of the orbital platform (inertia wheels, reaction thrusters, etc.).
Past structural systems have utilized metal, composite, or honeycomb core materials to achieve light, rigid structures, but have had difficulty in maintaining stringent alignment criteria at the same time. Thermal stability is a major limiting factor on many materials. Some composites and a few metals exhibit acceptable thermal properties, but unless they are fabricated and assembled into a structural system that takes advantage of these properties, the resultant assembly does not meet its accuracy or stability requirements.